This invention relates generally to gas turbine engine turbines and more particularly to methods for cooling turbine nozzles of such engines.
A gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure or gas generator turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. In a turbojet or turbofan engine, the core exhaust gas is directed through an exhaust nozzle to generate thrust. A turboshaft engine uses a low pressure or “work” turbine downstream of the core to extract energy from the primary flow to drive a shaft or other mechanical load.
The gas generator turbine includes annular arrays of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship. These components operate in an extremely high temperature environment, and must be cooled by air flow to ensure adequate service life. Typically, the air used for cooling is extracted from one or more points in the compressor. These bleed flows represent a loss of net work output and/or thrust to the thermodynamic cycle. They increase specific fuel consumption (SFC) and are generally to be avoided as much as possible.
Various methods are known for cooling turbine components including film cooling, internal convection, and impingement. Impingement is known to be a particularly effective cooling method and is frequently used in large turbine engines where the engine core flow is substantial. However, higher turbine stages in small turboshaft and turboprop engines stage do not typically employ impingement cooling of the airfoil because there is either not enough cooling air or enough supply pressure available. Instead internal features like turbulators or pins provide the necessary convection heat transfer enhancements.